The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
Among the engine components, relatively high temperatures are observed in the combustor section such that cooling airflow is provided to meet desired service life requirements. The combustor section typically includes a combustion chamber formed by inner and outer wall assemblies. Each wall assembly includes a support shell lined with heat shields often referred to as liner panels. In certain combustion architectures, dilution passages direct airflow to condition air within the combustion chamber.
In addition to the dilution passages, the shells may have relatively small air impingement passages to direct cooling air to impingement cavities between the support shell and the liner panels. This cooling air exits numerous effusion passages through the liner panels to effusion cool the passages and film cool a hot side of the liner panels to reduce direct exposure to the combustion gases. To facilitate cooling efficiency, each effusion passage is strategically located and formed with compound angles, e.g., a spanwise clock angle, i.e., azimuth and a surface angle, i.e., elevation. With lower emission and higher combustor operational temperature requirements, benefits accrue in response to efficient utilization of cooling air.